1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled stator vane with impingement and film cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases are discharged through a first stage high pressure turbine nozzle having stator vanes which direct the gases toward a row of turbine rotor blades extending radially outwardly from a supporting disk.
The turbine blades extract energy from the combustion gases and power the compressor. The gases are then channeled to a low pressure turbine typically having several stages of nozzle vanes and rotor blades which extract additional energy from the gases for producing output work such as powering a fan in a turbofan aircraft engine embodiment.
Since the high pressure turbine nozzle firstly receives the combustion gases from the combustor, it must be cooled for enjoying a suitable useful life. A typical turbine nozzle includes a row of airfoil vanes circumferentially spaced apart from each other and extending radially in span between outer and inner annular bands. The vanes are hollow for receiving therein a portion of compressor discharge air used for cooling the individual vanes. Especially in an industrial gas turbine engine, the first stage of stator vanes are exposed to the combustor outlet hot gas flow directly and therefore must be cooling the most of any of the turbine airfoils. The compressed air used to cool the first stage stator vanes is from the outlet of the compressor since the outlet air pressure is the highest pressure available from the compressor. The hot gas flow from the combustor has a pressure slightly less than the compressor outlet pressure due to losses from passing through the combustor. In order to prevent the hot gas flow ingestion into the film cooling holes of the stator vane, the cooling air pressure must be higher than the gas pressure. Because the high cooling amount required for the first stage stator vanes, the vane must be cooled with both impingement cooling and film cooling to withstand the high gas temperature. It is therefore a major design problem to be able to provide both impingement cooling and film cooling to the first stage stator vane with enough cooling air pressure so that the hot gas does not flow into the cooling holes.
In order to protect the external surface of the vanes from the hot combustion gases flowing thereover, various radial rows of film cooling holes are provided through the pressure and suction sides of the vane. Since the leading edge of the vane first receives the hot combustion gases, it typically includes several rows of film cooling holes in a showerhead configuration. The air discharged from the film cooling holes produces a boundary layer of cooling air along the external surface of the vane which is re-energized with additional cooling air from row-to-row. The film cooling air provides a barrier protecting the metal of the vane from the hot combustion gases during operation.
Since the combustion gases flow with different velocities over the pressure and suction sidewalls of the vane, the various regions of the vane from leading to trailing edge are subject to different amounts of heating therefrom, and correspondingly require different amounts of cooling. Since any air diverted from the combustor for cooling the nozzle vanes decreases overall engine efficiency, the amount thereof should be minimized while obtaining a suitable useful life for the nozzle vanes.
The varying heating effect of the combustion gases and the varying cooling effect of the cooling air further complicate vane design since temperature gradients are created. Temperature gradients cause differential expansion and contraction of the vane material, which in turn causes thermally induced strain and stress which affects the low cycle fatigue life of the vane during operation.
Several methods exist to improve the gas turbine engine power output and cycle efficiency. Lower the combustor pressure loss will yield higher turbine inlet pressure which results in a higher overall working pressure ratio across the gas turbine engine and a higher engine power output. A typical pressure loss differential through the combustor is 2.5 to 4.5% of the combustor pressure. Another method is to increase the turbine inlet temperature. Another is to reduce the total cooling and leakage flow demand. Less flow bleed off from the compressor for the turbine cooling results in more flow through the combustor to provide working fluid to the turbine. Another method is to minimize the cooling air pressure loss to the vane.
In the method of low cooling air supply pressure for the first stage vane cooling design, due to high loss of cooling air to the vane, the turbine inlet pressure is at a much higher level than the cooling supply pressure which limits utilizing effective film cooling in the cooling design for the vane. The first stage vane is then cooled by means of backside impingement cooling without an external showerhead film cooling. However, there is a limit of cooling effectiveness level that can be achieved for an all convectively cooled airfoil which translates to a maximum permissible operating temperature for the first stage vane material capability. On the other hand, in the case of low combustor pressure drop for the first stage vane cooling design, there is not sufficient cooling pressure to utilize both backside impingement cooling in combination with a showerhead cooling for the vane leading edge region. As a result of this design constraint, the turbine firing temperature and the potential growth for the engine is limited.
U.S. Pat. No. 6,200,087 B1 issued to Tung et al on Mar. 13, 2001 and entitled PRESSURE COMPENSATED TURBINE NOZZLE discloses a turbine vane with an embodiment in FIG. 4 of this Tung patent in which a leading edge cone shaped cavity is separated from a mid chord cavity by a slanted rib such that the two cavities are cone shaped in the cross section. Each cavity includes an impingement tube. Major differences between the Tung patent and the present invention is that the impingement tubes in the Tung patent do not extend out from the cavity far enough to solve the problem with the vena contractor, and the mid chord cavity is not directly connected to the outlet holes along the trailing edge. In the Tung patent, cooling air from the mid chord cavity flows through the impingement holes in the tube, then through the holes in the rib and through pin fins, and then through the outlet holes. Because of the extra row of holes in the rib, a higher pressure is required to prevent injection of hot gas into the outlet holes.
It is an object of the present invention to provide for a first stage turbine stator vane with both backside impingement cooling in combination with a showerhead cooling for the leading edge region of the vane.
It is another object of the present invention to provide for a gas turbine engine with a higher permissible operating temperature for the first stage stator vane.
It is another object of the present invention to provide for a process for operating an industrial gas turbine engine in which the overall back flow margin is within a minimum amount in order to allow for the first stage vanes to be cooled both by impingement and film cooling.
It is another object of the present invention to produce a first stage turbine stator vane with both impingement and film cooling in which the pressure drop across the impingement cooling holes is no more than 30% of the total pressure drop across the impingement and the film cooling holes of the leading edge.